The present invention relates to solid rocket motors. In particular, the invention relates to solid rocket motors which do not require the use of an expansion nozzle to control the exhaust gases.
Solid rocket motors operate by burning a solidified mixture comprising a fuel and an oxidizer, producing a large volume of gaseous combustion by-products. These combustion products are vented from the rocket motor at high speed, producing thrust in the opposite direction from which they escape the motor. This thrust is used to accelerate the vehicle. For example, solid rocket motor boosters are used to help accelerate the Space Shuttle at launch in order to boost the orbiter to an altitude and speed from which the main engines alone can propel it into orbit.
In typical rocket vehicles, the payload represents a very small fraction of the overall vehicle mass at launch, most of the mass consisting of propellant and engine structure. Because any required engine structure must be lifted into orbit along with the intended payload, any increase in engine structure mass requires an increase in the amount of propellant used to launch the vehicle. Engine structure is essentially payload that the engines must lift.
Because of this, a rapid increase in the amount of fuel required, and hence the overall rocket thrust required, occurs as engine mass increases. A rocket with the lightest possible structure is desirable, so that the maximum fraction of the total vehicle will consist of useful payload and propellant, rather than supporting structures. These structures, such as engines and mechanical connections, are essentially deadweight.
Despite this desire to minimize the mass of support structure in the rocket vehicle, the nature of rocket propulsion generally requires that certain structures are used in order for the rocket motor to properly function. Specifically, high internal temperatures and pressures are needed for proper combustion and efficient thrust in a typical solid rocket motor. Designing the vehicle to withstand these extremes results in additional structure, which adds mass to the rocket and reduces the payload fraction.
To produce thrust, combustion by-products are exhausted from a rocket motor through an aperture at the rear of the motor which opens into a channel. This channel is wider farther from the aperture and opens directly to the ambient environment at its farthest point from the aperture. This expanding channel is referred to as the xe2x80x9cnozzlexe2x80x9d of the motor. Often, the aperture is considered part of the nozzle as well.
A typical nozzle must be able to withstand not only the pressures to which it will be subjected by the combustion products of the engine, but also must withstand the high temperatures and corrosive nature of the exhaust gas flow through it. Additional aerodynamic stresses are imposed by the nozzle""s passage through the atmosphere. As a result, a nozzle which can operate under these conditions often adds significant mass and complexity to a rocket""s structural design. This in turn requires greater thrust, and hence a larger and more costly rocket system.
In a typical solid rocket motor, the aperture constricts the flow of the exhaust gas as it passes from the rocket motor into the expansion channel. This constriction increases the pressure on the exhaust gases, and consequently, increases the pressure within the rocket motor itself. This additional pressure caused by constriction on the exhaust gases is known as xe2x80x9cback-pressurexe2x80x9d. Although increasing the internal pressure in the rocket requires a corresponding increase in the structural strength of the rocket motor casing, it is traditionally desirable to do so anyway. This is because the solid fuel used in a traditional rocket motor will only burn properly at pressures much higher than the ambient pressure. Without the constricting aperture providing additional back-pressure on the rocket motor, the traditional fuel grain would not burn with enough intensity to produce the desired thrust, and the fuel would be wasted. Introducing a constricting aperture into the exhaust flow path raises the pressure of the exhaust gas and provides the necessary back-pressure to ensure a fast, effective burn of the fuel.
While the aperture is used to control the pressure and expansion of the exhaust gas inside the rocket motor, the nozzle of a typical rocket motor is used to control the expansion and pressure of the exhaust gas as it leaves the rocket motor. Such control over the expansion rate of the exhaust gases is needed because traditional rockets lose much of their efficiency and thrust if their exhaust gases are allowed to vent in a turbulent manner. Ordinarily, turbulence will result when the pressure of the exhaust gas is significantly different from the ambient pressure into which the gas is vented. By expanding the exhaust gases and reducing the pressure in these gases, a nozzle minimizes the turbulence in the exhaust and increases the thrust. In the absence of a nozzle, the efficiency of a traditional motor drops to the point where the thrust produced is undesirably low.
Although it would be desirable to eliminate the additional mass and complexity of a nozzle on a rocket engine, traditional designs do not produce enough thrust to make such a design feasible for lifting a payload without the use of an expanding nozzle and constricting aperture.
In the present invention, a high burn-rate solid propellant matrix is used, eliminating the need for additional pressure inside the rocket motor and also eliminating the need to minimize exit turbulence in the exhaust gas flow. By eliminating the nozzle and aperture, mass reductions are made in the structure of the rocket, which allow for a greater payload fraction and greater mass efficiency of the rocket. The simplifications of design that are made possible also allow for more flexibility in structural design of the rocket.
In a preferred embodiment of the present invention, a rocket casing is attached to a payload. The casing is filled with a solid propellant matrix and contains at least one opening which provides an exhaust path for any combustion by-products to exhaust directly from the interior of the rocket casing to the ambient environment. When operating, the propellant matrix is ignited and burns, producing exhaust gases that vent directly through the exhaust opening, producing thrust in the opposite direction.
Additionally, the propellant matrix may be comprised of a solid homogeneous mixture of fuel particles that are distributed within a matrix of solidified oxidizer. The propellant matrix may also comprise an intimate stoichiometric mixture of oxidizer and metallic fuel particles, or it may comprise a substantially homogeneous mixture of metallic fuel particles embedded in a matrix of solid oxidizer wherein the average distance between the metallic fuel particles is controlled.
The propellant matrix may also be formed such that the burning surface is initially located at the lower, or exhaust end of the rocket, and progresses as it burns toward the top, or payload end, of the rocket.
In a further preferred embodiment, the rocket casing comprises a consumable material, such that the casing will burn away as the propellant matrix is consumed and the casing is exposed to the heat and pressures of the exhaust gases.
In another preferred embodiment, a solid rocket motor comprises two solid propellant matrices, one inside the other. The outer matrix is stiffened and formed into a casing for the inner matrix, and is attached to the payload to be accelerated. The inner fuel matrix is designed to have a higher burn rate than the outer fuel matrix, so that the inner matrix remains contained within the outer matrix throughout the operation of the motor.
Another preferred embodiment of the present invention is a system wherein solid fueled booster rockets are used to accelerate a payload delivery means, and each booster makes use of a solid propellant matrix within a casing and a direct opening from the interior of the rocket casing to the ambient environment. Such boosters are attached to the payload delivery means and provide thrust for said payload delivery until they burn out and are jettisoned.